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Abstract

The large, EU Supported ESPOSA (Efficient Systems and pro-pulsion for Small Aircraft) project has developed new small gas turbines for small aircraft. One of the important tasks was the engine - airframe aero-thermal radiation integration that included task of minimizing the infrared radiation of the small aircraft, too. This paper discusses the factors influencing on the aircraft infrared radiation, its possible simulation and measure-ments and introduces the results of small aircraft infrared radia-tion measurements. The temperature of aircraft hot parts heated by engines were determined for validation of methodology devel-oped and applied to engine - aircraft thermal integration.

Keywords

aircraft, infrared radiation, radiation measurements, engine - airframe thermal integration

1 Introduction

The infrared is an invisible radiant energy. This electromag-netic radiation has longer wavelength (starting from nominal red visible line wavelength 700 nanometers up to 1 mm), than

the visible light. The first defense equipment for measuring the

infrared radiation were developed prior the World War I. During World War II, the infrared detection was used for tracking, too

(Hudson, 1969). The first missiles seeking the IR, i.e. infrared

(thermal) radiation had been developed for beginning of 1950s, while, because the success of radar systems, its wide applica-tion was started about 10 - 15 years later. Titterton (2006) found from statistics that, the heat-seeking missiles had been respon-sible for more than 80 % of all the combat losses over 40 years

since 1960s. During the Gulf war, (1991) IR (infrared) missiles

were responsible for 76% of the neutralized aircrafts, while during the Kosovo war, the NATO aircraft were not allowed to

fly below 15 000 ft, because the missiles effectiveness (Santos

et al., 2007). On the same time, since 1987, for 26 years 35

civilian aircraft had been attacked by shoulder-fired missiles

that caused more than 500 deaths (Bolkcom et al., 2004). (It is true, this large number might be less, because some lost civilian aircraft were used for military purposes, some others or

bet-ter to say most of them were attacked in conflict zones, where

they may have been perceived as being used for military

pur-poses (Sweetman, 2003)). In any case, in 2003, Department of Homeland Security initiated – probably the first – large civil -ian project for detailing the implementation of an anti-missile system on a single commercial aircraft (Bolkcom et al., 2004).

The aircraft infrared radiation has investigated by many researchers, developers for the last 50 - 60 years as it summa-rized by Mahulikar et al. (2007). The most of research were and

are related to military applications. For first period, the experi -mental studies had been focused on the military aircraft infrared signature measurement and aircraft type recognition, too. Later, the practical measurement of the civilian large aircraft infrared radiation signatures (Coiro et al., 2012), the radiation modeling and countermeasure became to set of important tasks. Parallel,

to these, the infrared radiation sensor technology (Rogalski,

2012) has been developed rapidly, too. The information about 1 Department of Aeronautics, Naval Architecture and Railway Vehicles,

Faculty of Transportation Engineering and Vehicle Engineering, Budapest University of Technology and Economics,

H-1521 Budapest, P.O.B. 91, Hungary

2 Aviation Training Centre,

Rzeszow University of Technology,

al. Powstancow Warszawy 12, 35-959 Rzeszow, Poland

3 CIRA Centro Italiano Ricerche Aerospaziali,

Via Maiorise 81043 Capua (CE), Italy

* Corresponding author, e-mail: jrohacs@vrht.bme.hu

https://doi.org/10.3311/PPtr.11514

Creative Commons Attribution b

research article

PP

Periodica Polytechnica

Transportation Engineering

Small Aircraft Infrared Radiation

Measurements Supporting the Engine

Airframe Aero-thermal Integration

Jozsef Rohacs

1*

, Istvan Jankovics

1

, Istvan Gal

1

, Jerzy Bakunowicz

2

,

Giuseppe Mingione

3

, Antonio Carozza

3

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errors and uncertainties in the IR measurements could be found

in Minkina et al. (2009).

The research can be divided into two major groups: inves-tigation of the aircraft infrared signature (including the theo-retical and practical studies) and sensing the infrared radiation.

Both have the same objective target recognition. Several papers

list the codes using in investigation of the aircraft infrared radi-ation (Coiro et al., 2012; McGlynn, Auerbach, 1997; Mahulikar

et al., 2007) from SIGGE of Swedish Defence Research that

can be applied to compute the aircraft infrared signature,

espe-cially effects from nozzles and plumes; trough the US code SPIRITS, NATO’s NIRATAM validated on large set of military aircraft; code SIRUS generated by large companies Lockheed Martin UK INSYS and BAE Systems until commercial codes SE-WORKBENCH, Radtherm-IR, SAFIR and SEISM. The

Belgium Defence establishment has developed an open-source

code OSMOSIS for compute the ships infrared signature that

can be applied to the Arial vehicles, too.

Nowadays, the small and personal air transportation system

developing rapidly (Holmes et al., 2004; Piwek and Wiśniowski, 2016; Rohacs, 2007a; 2007b). The ESPOSA (2016) project had

developed special small gas turbines for small aircraft. During optimization of the engine components and airframe integra-tion the thermal condiintegra-tions and the infrared radiaintegra-tion had been taken into account (Buonomo et al., 2013; Carozza et al., 2013;

Hendrick et al., 2015; Veress, Bicsak and Rohacs, 2016). The helicopters flying at low altitude also could be treated by IR missiles (see for example (Jingzhou et al., 2014)).

This paper shows and discusses the results of the infrared radiation measurements.

2 The sources of the aircraft infrared radiation

The infrared radiation of aircraft is composed from the dif-ferent sources (Fig. 1) including the radiation from the aircraft

hot parts, radiation from the heated parts, reflections of the

sun-, sky- and earthshines.

The infrared radiation like all electromagnetic radiation interacts with matter in a variety of ways (White, 2012):

• reflects - a wave from a surface (angle of reflection equals

the angle of incidence),

• refracts - when the direction of a wave bends when pass-ing between two transparent media with different

propa-gation speeds (Snell’s law),

• scatters - that occurs (in blue sky) upon interaction with particles whose size approaches the length of the wave, • diffracts - interaction occurring around the edges of

an obstruction,

• interferes - interaction in both a constructive and destruc-tive manner,

• absorbs - when radiation is converted into another form of energy (mostly to heat),

• emits - from matter by conversion from another form of energy,

• transmits - propagation of the infrared radiation through a transparent medium (or vacuum),

• polarizes - when an electric field is partially polarized by reflection from dielectric.

Fig. 1 Sources of aircraft infrared radiation (created by use of (Mahulikar et al., 2007)).

3 Governing laws

The aircraft infrared radiation that can be measured or sensed depends on heat sources, and intensity of sources (initiated by

fuel burning (Bicsák, Hornyák and Veress)) aircraft real flight situation, inside and outside flow conditions and the infrared radiation itself. The last one is defined by the following - gov

-erning - rules (Siegel and Howell, 1972).

3.1 Planck’s radiation law

It describes spectral radiance (L - energy per source unit parameter) that is emitted by a black body in thermal equilib

-rium at the defined (absolute) temperature (T ). It can be given

in different forms. For example the spectral radiance per unit of

wavelength (λ) equals to

L T L hc

e

e hc kT

λ

λ

λ λ

, , , / ,

(

)

= ⋅

2 1

1

2

5

where k – Boltzmann constant: 1.38 × 10-23 J/K, h – Planck

constant: 6.626 × 10-34 J·s, c – speed of light: 299 792 458 m/s.

The SI dimension of the L(λ,T) is W·sr-1·m-3. (Here sr is the SI

unit of the steradian or square radian angle, Ω.).

3.2 Wien’s displacement law

It defines the maximum wavelength (λmax) at which the spec-tral radiance of black body radiation per unit wavelength is

maximum, too.

λmax = , max = . ⋅ ⋅ ,

b

T E 1 309 10 T

7 5

where b – Wien’s displacement constant,

b= × − m K

2 8977721 103

. .

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[image:2.595.314.555.139.280.2]
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3.3 Stefan Boltzmann radiation law

Integration the Planck’s distribution on the whole range of

wavelengths (λ = 0 ÷ ∞) results to total energy emitted by a

black body per unit surface area per unit time.

Q=∞

E

(

T d

)

= ⋅T

0

4 λ, λ σ ,

where σ is the Stefan-Boltzmann constant, σ =5 67 10⋅ −8

(

2 4

)

. W m K .

3.4 Black body

A black body is defined as an ideal body that allows all the

incident radiation to pass into it and absorbs internally all the

incident radiation. This means no reflected and transmitted energy exist. This is true of radiation all wavelengths and for all

angles of incidence. Hence the blackbody is a perfect absorber

in incident radiation (Siegel and Howell, 1972).

Blackbody is an ideal of comparison for a body emitting radiation. A blackbody has two main properties in thermal

equilibrium. First, it is an ideal emitter at every wavelength, it

emits as much or more energy than any other body at the same

temperature. Second, it is a diffuse emitter, which means that

the energy is radiated isotopically, independent of direction. The real body radiation can be given by its emittance or emissivity that is a ratio of real and black body radiations:

ε = E Eblackreal .

The emissivity varies according to the surface properties, the material and, for some materials, also according to the tem-perature of the object. It also depends on the wavelength and angle of incidence.

For rough surfaces the value is close to unity, for polished surfaces it is around 0.02.

3.5 Kirchhoff’s law of thermal radiation

The electromagnetic infrared radiation as usually hits the objects, while the objects absorb some of this energy. The value of absorptivity (α) is between 0 and 1, where 1 means a black body, 0 means a perfect white body.

The heated body, having higher temperature emits more infrared radiation than colder objects.

It is easy to understand that, in case of thermal equilibrium,

when the temperature of a body is constant, i.e. independent on

time, the emitted and absorbed radiation must be equal. Practically

this means the body gives off as much energy as much absorbed: ε α= .

3.6 Conservation of radiation energy

When the incident radiation reaches the surface of a material, it can act 3 ways depending on the material properties: it can be

absorbed, transmitted and reflected. The absorbance (α) describes

how much fraction of incidence radiation is absorbed, the trans-mittance (τ) is a measure of the ability of a material to transmit

(allow through) infrared radiation. The reflectance (σ) is a

mea-sure of the ability of a material to reflect infrared radiation.

By use of energy conservation law the sum of absorbed,

transmitted and reflected energy must be equal to the incident:

   

Q= ⋅ + ⋅ + ⋅α Q τ Q σ Q,

or

α τ σ+ + =1.

As transmission rarely plays a role in practice, the transmis-sion (τ) is omitted and the formula is simplified to

α σ+ =1

or because of Kirchhoff’s law ε σ+ =1.

This last equation allows determining the emissivity of the materials from its reflection that is easier to measure.

This short summary of governing rules contains all the

required "radiation" equations applying in aircraft infrared

radiation simulation, measurements and countermeasures. The other models are given by special studies and papers. For

example, the basic equations of the fluid mechanics are sum

-marized in (Buonomo et al., 2013; Veress, Bicsak and Rohacs, 2016). One of the recent examples of in-flight testing concern -ing measurement of boundary layer by infrared means might be found in Boden et al. (2015).

4 Practical aspects

According to the knowledge on (i) material properties, (ii)

user manuals and guides of infrared videos (see for example

(Pocket, 2012)) and (iii) our practice with developing the ther-mographic evaluation methods the following important practi-cal aspects must be underlined.

There are some aspects according to the emittance: (i) for

real bodies ε < 1, because real bodies also reflect and possibly transmit radiation; (ii) many non-metallic materials (e.g. PVC,

concrete, organic substances) have high emissivity in the long-wave infrared range that is not dependent on the temperature

(ε ≈ 0.8 – 0.95); (iii) metals, particularly those with a shiny sur

-face, have low emissivity that fluctuates with the temperature;

(iv) ε must be set properly for accurate measurement.

Another important comments on reflectance are (i) σ

depends on the surface properties, the temperature and the type

of material; (ii) in general, smooth, polished surfaces reflect

more strongly than rough, matt surfaces made of the same

material; (iii) the temperature of the reflected radiation must be taken into account (Reflective Temperature Compensation - RTC) that in many measurement applications corresponds to

the ambient temperature (mainly in indoor thermography); (iv) (3)

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the RTC can be determined using a Lambert radiator (when the incident radiation is reflected with equal strength in all direc

-tions like the black bodies) and (v) the angle of reflection of the reflected infrared radiation is always the same as the angle of incidence, (similar to the reflection in optic).

Other comments according to the transmittance are (i) dependence on the type and thickness of the materials and (ii) the fact, the most materials are not transmissive, i.e. permeable, to long-wave infrared radiation.

Generally, there are two types of reflection can be distin

-guished, the specular and the diffuse reflections.

Specular reflection is a mirror-like reflection of electromag

-netic waves (Fig. 2). In the optics the specular reflection means that light reflects off surfaces in a very predictable manner in accordance with the law of reflection: angles between the sur

-face normal and incident and normal and reflected radiation

are the same. In thermography it means that the heat source in the background is clearly visible on the surface because of

specular reflection. The other result of this kind of behavior that the reflected picture of radiation sources highly depends on the angle of view. By changing the angle of view the reflected

background can be optimized.

Fig. 2 CFM56 engine exhaust reflection on the wing bottom surface

Diffuse reflection is a reflection of rough surfaces. The

roughness of the surface means that each individual ray of inci-dent radiation reaches a surface which has different orienta-tion. The normal lines at each point of incidence of rays are

different for different rays. So the reflected radiation is scatter -ing depend-ing on the surface roughness. It means, with given accuracy of measurement of incident radiation, the error of measurement of emitted energy will be proportionately higher

for reflective surfaces. Accurate measurement requires precise

determination of emissivity values.

Finally, the Lambert’s cosine law says that the radiant intensity observed from an ideal diffusely reflecting surface is

directly proportional to the cosine of the angle between the sur-face normal and the direction of observer.

The emissivity may vary from 0.02 (aluminum, not oxidized

at 25 °C) up to 0.94 (glass - 90 °C).

The accuracy of the thermo radiation measurements depends on the environmental conditions, too. The clouds, precipitation,

air humidity, air flows, air pollution, etc.

The objectives of the thermo radiation measurements are the

evaluation of influence of aircraft structural solutions and their

possible optimization for reducing the infrared radiation. Our practice going back to the end of the last century (Oravecz et al., 2000), and it showed that even relatively simple infrared cameras (like AGA Thermovision 750 type working with a

spe-cial video recorder used in first investigations) can be applied

to measurement, analysis and evaluation of the aircraft infrared

radiations. As example, the Fig. 3 demonstrates the changes in infrared signature of an military helicopter after equipping with

a special thermal emittance reducer. The photos were made in

1998, at the same flight conditions, from same distance, about

the same military helicopter without (upper photo) and with

(lower picture) equipment for IR reduction. The photos demon -strate that, the infrared radiations origins from all the heated

parts of helicopters. The left photo shows the radiation reflected

from the rotor blades that origins from the helicopter itself.

Fig. 3 Infrared radiation of a military helicopter without (upper) and with the thermal emittance reducer (lower image)

During preparation for measuring the infrared radiation

of ESPOSA (2016) small demonstration aircraft, there were

[image:4.595.343.529.335.640.2] [image:4.595.68.268.364.516.2]
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The Table 1 contains explanation and pictures of infrared

radiation of engine nacelles and plumes.

The Fig. 4 and 5 show the changes in temperature of nacelle surface during the engine tests. As it can be seen, the tempera-tures measured at the M2 and M3 points are correlated by the

N1 engine rotation speed (that - in Fig. 4 - given as % of maxi

[image:5.595.299.546.114.488.2]

-mum - take off – speed).

Table 1 Engine nacelle and plum radiation

Engine at low RPM (engine rotation speed - revolution per minute), the exhaust plume on thermal picture is short, with moderate temperature.

Engine during RPM increase, the edge of exhaust is smooth while its length has increased, and relatively high temperatures can be observed in the core.

Engine at high RPM, length of exhaust is long, edges are sharp, temperatures in the core are higher, but not as high as during the transition.

Fig. 4 Points in which the nacelle surface temperatures are measured (in time interval with no measured data, the camera was moved to make measurements

from different point of view)

The Fig. 5 shows eight thermo images of the nacelle that were selected with nearly same time intervals, to show how changing the temperature of engine nacelle during the 1100 seconds long time period.

Fig. 5 Infrared pictures of an engine nacelle at engine test (time is sec after starting the test)

5 Radiation simulation

The radiation simulation is a rather complex task, because it depends on the internal/external heat sources (as gas turbine, exhaust gas, oil radiators, hydraulic actuators, electric equipment, sunshine, etc.), internal/external flows (engine-airframe integra

-tion), infrared radiation, ray reflections and so on. In open and

available references all these aspects are investigated.

One of the first simulation model of the aircraft infrared sig

-nature, the NIRATAM (the NATO Infrared Air Target Model) was developed by the NATO AC 243, Panel IV, Research Study Group 6 (RSG-6), for beginning of 1990s’ (Noah et al., 1991). It was based on theoretical studies, field measurements, and infrared

data analysis performed over many years. The vehicle fuselage, facet, model includes radiation due to aerodynamic heating,

inter-nal heat sources, reflected sky, earth, and solar radiation. Plume

[image:5.595.45.337.163.722.2]
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generated graphical outputs of the target wireframe, plume flow field, atmospheric transmission, total signature and plume signa -ture. This methodology is well applicable today, too.

The early 1990 several general models were developed.

For example, the SPIRITS as spectral and in-band radiometric

imaging of targets and scenes in a modeling software could be applied to model targets of a generic aircraft (Miller, 1993).

Already in 1998, Joyner et al. (1998) and his colleague

cre-ated a generalized model, namely they integrcre-ated three IR mod

-els: Spectral In-Band Radiance of Targets and Scenes (SPIRITS), Physically Reasonable IR Signature Model (PRISM), and IR Electro-optical Naval Engagement (IRENE) into sensor simulator.

For the beginning of the new century, the complex mod -eling using the CFD results as inputs had been developed (Andersson, 2002).

a) The calculation of the aircraft infrared radiation and sig-nature simulation depends on the objectives. The most

used objective is the simulation of the IR signature from the recognition point of view (Santos et al., 2007; Willers

and Willers, 2012). According to this, there are several important key considerations and aspects determining

the creation of the required models and the simulation

methodologies should be taken into account.

The key considerations identified by Willers and Willers

(2012) for such an imaging infrared simulation system are: (i) radiometric accuracy in all spectral bands, (ii) accurate emitting source surface temperature behavior,

(independently on the heat sources), (iii) high fidelity geo

-metrical and spatial texture modeling, (iv) true dynamics and kinematic behavior (in six degrees of freedom), (v)

detailed modeling of signatures and backgrounds, (vi)) accurate atmospheric transmittance and path radiance models, (vii) realistic rendering of the scene image in radiometric, spatial and temporal terms, and (viii) com-prehensive sensor modeling to account for primary and second order imaging effects.

These key factors determine the requirements to develop the simulation models. The following aspects define the required modeling approaches from the (i) image record -ing, (ii) environmental conditions (as humidity) (iii),

air-craft flight condition (spatial orientation), (iv) radiance and (v) reflection simulation point of view.

b) Sensor work. The sensors collect images for each wave number, η (η = 1/λ), therefore the radiation must be deter -mined for all spectral band.

c) Sensing feature. The sensors record series of rays sepa

-rated from the neighboring rays with angles of ∆θ and ∆ϕ

(see Fig. 6.a).

d) Distance effect. Generally, the receivers sense the

radi-ant flux (Φe radiant energy emitted, reflected transmit -ted or received per unit time) in form of radiant inten-sity (I – flux per steradian angle) and/or irradiance (i.e.

area power density at the receiver, E – flux per unit area). Irradiance and intensity are related by the square of the

distance: E = I/R2.

e) Projected area of source. According to the previously

described definitions and aspects, the sensor elements "look at" the areas as spherical areas defined by the stera -dian angles, Ω (Fig. 6.b). That means the projected areas

might be defined to all the aircraft surface elements as the

solid body facet model elements that will be perpendicu-lar to lines connecting them to the sensor elements. The angle between the normal to sensor plane and

connect-ing line, θreal and spatial position of the surface elements

define the projection areas, Aproj.

f) Aircraft position. The measurable radiance depends on

real position of aircraft. At first the heated elements of the aircraft can be "seen" by receivers directly, or partly, or even the heated elements can be in "shadow"of other

aircraft elements. On the other hand, it is easy to

under-stand the reflection of the sunshine, skyshine and earth -shine really depends on the position of aircraft related to the sensor position.

g) Atmospheric filtration effects. The molecules and par

-ticles present in the atmosphere filter out some part of thermal radiation. For example, the infrared radiation generated by the fighter tailpipe and the plume is mostly

contained in range of wavelength 3 to 5µm. The

measure-ment (Santos et al., 2007) shows that, the radiant spectral

emittance in the given range of wavelength is nearly zero

with two large peaks (about 40 – 70 times greater values)

in range between 4.17 - 4.20 µm and 4.32 - 4.70 µm. h) Absorptivity. The atmospheric filtration effect can be deter

-mined by use of absorptivity of gases. The spectral trans-missivity and absorptivity in a homogenous isothermal gas

cell can be given by use of absorption coefficient, Kη, τη = η αη = − η

− −

e K s e K s

, 1 ,

where s is a thickness of the gas cell. Because the

absorp-tion coefficient depends on the contents of the gas cells, therefore the absorption coefficient is the sum of coeffi

-cients defining the gas cells with given concentration of the

elements as CO2, H2O, etc. (Andersson, 2002). The coeffi

-cients might be determined from the (Rothman, 2013).

i) Modeling aspects. Actually, the aircraft infrared signature measuring is based on the radiant intensity or irradiance measurements (see sub point c)) that are analogous to radar

cross section (RCS) detection or simulation (see for exam

-ple (Rohacs, Palme and Sket, 2010; Palme et al., 2010;

Perotoni and Andrade, 2011)). However there is a

princi-pal difference in analogies between IR and radio frequency (RF) because the target aircraft in the IR is an active emitter rather than the passive reflector of a distant RF illuminator.

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RF effective radiated power (ERP), which combines trans -mitter power with antenna beam (White, 2012).

j) Complexity. There are several methodological approaches

to determining the infrared signature of aircraft. The

com-plexity of the problem, simulation the infrared signature, can be characterized by one of the first code developed by Aerodyne Research Inc. (USA) in 1989 already to simulate

long-range air-to-air detection and tracking engagements. The code integrated eight standalone modules (Iannarilli,

Wohlers, 1991): (i) SPIRITS (aircraft IR signatures imag -ing module), (ii) CLOUD (sky background imag-ing

mod-ule), (iii) LOWTRAN, (iv) TRACKER (signal processing and tracking module), (v) IPAS (optical sensor and spatial processing module), (vi) MISSION (dynamic trajectory module), (vii) ENGAGER (integrates all modules), and (viii) HIGH-LEVEL SCENARIO SPECIFIER (user-inter -face module).

k) Simulation aspects, inputs. The inputs often classified as

(Davis and Thomson, 2002) (i) background information (geography, date and time, meteorological inputs), (ii) plat-form characteristics (size, shape, materials, coating,

pro-pulsion and auxiliary equipment) and (iii) threats (wave -length band, spatial resolution as scanner, imaging, etc., and sensitivity including detector noise). Here the platform might be the critical elements, because the applied mate-rials and especial their coating and painting (Mahulikar et al., 2006) and fuselage skin heated by engine thermal pro-cesses. The last one can be determined by use of model developed by Mahulikar and his colleague (Mahulikar et al., 2001; 2005) that incorporates the radiation interchange in engine layout, to compute the temperature distribution of rear-fuselage skin for given engine operating conditions.

The third platform problem is associated by engine exhaust

plume that is well analyzed by (Mahulikar et al., 2007). Of

course, the exhaust plume radiation together with the real

fuselage play deterministic role in case of military aircraft and sensing from the real side. During operating the after burner system, the plume might be longer then the fuselage.

l) Methodology. Most of the simulation methods

com-bine the computation fluid mechanics thermal condition calculation radiation emission and radiation reflection

(Mahikular et al., 2001; Anderssen, 2002; Jianwei and Qiang, 2009; Pan et al., 2011).

The ESPOSA (2016) project studied the infrared radiation

of small aircraft by use of two major methodologies. The theo-retical investigations had objectives aero-thermal analysis and evaluation of the engine airframe integration and reducing the temperature distribution (by this infrared radiation) of the air-frame elements heated by working engines (Buonomo et al.,

2013; Carozza et al., 2015; Veress, Bicsak and Rohacs, 2016).

The other program was the practical measurements of the

infrared radiation of engine – airframe sections of the aircraft demonstrators equipped by small gas turbine developed by ESPOSA project.

The aircraft infrared simulation is based on the coupled

multiphisical (numerical fluid dynamics - CFD and heat trans -fer simulation) and / or multi-disciplinary models (including

the chemical equations defining the burning processes and contents of the exhaust gases). Fig. 7 shows the applied simu -lation methodology.

The applied computational fluid dynamics simulations couple flow and thermal fields, solving the thermal conduction inside the

nacelle wall (conjugate heat transfer analysis) due to the pres-ence of the engine hot components and taking into account for nacelle cooling/ventilation system (Buonomo et al., 2013). The CFD analysis with proper boundary conditions at engine surface and nacelle wall considers the both convective and radiative heat transfer. Three-dimensional CFD calculations were performed by

solving the Reynolds-averaged Navier-Stokes (RANS) equations. By use of Fluent code, the density-based flow solver and the stan

-dard SST (Shear-Stress Transport) k-ω model for turbulence were employed. In addition, the radiative exchange between engine

and nacelle was also taken into account, by using the Discrete Ordinate model (Buonomo, 2013).

Fig. 6 Thermal radiation rays drawn from sensor point of view (a. (left side) - sensing the neighboring rays by steps of ∆θ and ∆ϕ, b. (right side) - the sterian angle, Ω and measured surface, Asurf and projection area, Aproj)

[image:7.595.62.543.43.192.2]
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Fig. 7 Aircraft infrared radiation simulation methodology

6 Theoretical investigations

The ESPOSA (2016) project had resulted to implementing

the developed small gas turbines on several small aircraft

dem-onstrators. One of them is the EM-11 ORKA aircraft designed and manufactured by Margański and Mysłowski. This type was

originally designed and built with two piston engine in pusher

configuration, but was redesigned later and equipped with two TP-100 turbine engine designed and manufactured by PBS Velká Bíteš. The aircraft is a high wing, 4 seated aircraft, with

retractable tricycle landing gear arrangement, and pusher

pro-peller configuration (image of this aircraft is used in Fig. 1).

The another aircraft was a low wing, 4 seated I-31T aircraft (Fig. 8) designed by the Institute of Aviation in Warsaw and

equipped with TP-100 turbine engine, too.

Fig. 8 The ESPOSA aircraft – demonstrator I-31T

According to the installation (2016) manual the maximum

allowed operational temperatures for the TP-100 are given in Fig. 9. The Figure shows the hot sections that heat the nacelle elements.

The Figs. 10 and 11 demonstrate the results of the theoretical investigations.

The Fig. 10 shows the static temperature distribution on

the nacelle surface of EM-11 ORKA aircraft (Buonomo et al., 2013) determined for cruise condition, namely for flight

with cruise speed 87.445 m/s at 2750 m. At this level the air

temperature equals to 268 K. The radiative heat transfer con

-tribution had been taken into account, too. Its influence is in expected areas of the nacelle close to the hottest part of the engine. Further, no influence is foreseen in the nacelle parts far

enough from engine and where forced convection effectiveness prevails, as the combustion air intake, at the lower front face

of nacelle. The maximum temperature value (about 380 K) is reached close to the engine pipes and gas generator zone. This

maximum temperature is close to maximum allowable one by

component materials of nacelle skin (Buonomo et al., 2013).

[image:8.595.43.296.49.283.2]

Fig. 9 Static temperature (K) contour of nacelle of EM-11 ORKA aircraft (Buonomo et al., 2013)

[image:8.595.310.560.251.476.2] [image:8.595.63.277.534.655.2] [image:8.595.322.542.538.713.2]
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The Fig. 11 shows the wall temperature contours of nacelle

of the same EM-11 ORKA aircraft for the engine idle condition

(Carozza et al., 2015). For easier evaluation of the results the idle condition was chosen as 101325 Pa air pressure and 300 K air temperature. The propeller effects had been taken into account. (Carozza et al., 2015). Generally, the propeller may reduce the skin temperature up to 35 degree (Carozza et al., 2015). This

fact explains the relatively low temperatures on nacelle surface.

As it can be recognized from Figs. 10 and 11 difference in

maximum nacelle wall temperature for cruise and idle condi

-tion reaches 100 degree. However, the zones of the maximum

[image:9.595.313.540.46.183.2]

temperatures are very small.

Fig. 11 Wall temperature (K) contours in case of engine idle condition and taking into the propeller effects (Carozza et al., 2015)

7 Measurements

ESPOSA aircraft demonstrators EM-11 OKRA and I-31 T

(see Figs. 1 and 8) were applied in practical measurements of the infrared radiation of engine - airframe sections. The engine ground tests were measured. A Testo 885-2 thermal imager was

used. This 320 x 240 pixel camera records the infrared radia

-tion of spectral range 7.5 – 14 μm, by image refresh rate 33 Hz. During measurements the engine parameters (frequency of rotation, exhaust gas temperature, oil temperature) and tem -peratures under nacelle by use of thermocouples (T1 - put close to the engine control unit, T2 - near the top of the engine cover, around the gas generator, near the inner surface of the mask,

T3 – near the generator over the gearbox and T4 – near the

fuel oil pump (see Fig. 12)) were registered, too, for supporting the further evaluation of the theoretical studies and practical measurements.

The quality of evaluation depends on the measuring situa -tion as posi-tion of sun, wind direc-tion, humidity, etc., too. The Fig. 13 shows the position of the investigated aircraft EM-11

OKRA during the test.

[image:9.595.308.541.208.458.2]

Fig. 12 Thermocouples under the nacelle

Fig. 13 Measurement condition

The Fig. 14 presents images recorded during the measure-ments. The left engine was started at 2:47 PM, i.e. at 46020

sec of the given day. The test was finished at 4:08 PM. the test

included the series of changes in operational condition as it

given in Table 2. So, the Fig. 14 shows images for engine start, taxiing, idling, off, max cruising, cruising, approach,

[image:9.595.46.280.231.439.2]

take-off, stopping engine regimes.

Table 2 Engine test process

Time Engine operational

condition Time

Engine operational condition

46020 Left engine start 49680 Approach

46200 Left engine warm up 49740 Take-off 46440 Right engine warm up 49860 Approach

46620 Taxiing 49920 Taxiing

46920 Idling 50220 Engines stop

47520 Take-off 50280 Cooling

[image:9.595.303.551.616.769.2]
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[image:10.595.46.295.51.452.2]

Fig. 14 Examples of infrared images recorded during an engine test

In Fig. 15, there is the same image showing the thermal con-dition of the left engine nacelle after engine test with use of two different temperature scales. As it can be seen, the hottest part of

nacelle has appeared around the engine exhaust tube (see Fig. 16). The maximum temperature may over 100 0C that is well correlat-ing with the theoretical calculations (see Figure 10). The

investi-gated engine is a turboprop engine; therefore the exhaust plume has not considerable thermal radiation. The left end of the exhaust

pipe has top radiation, because the infrared camera looks into the pipe (as it can be understood from comparison the pictures of

Fig. 15 and Fig. 16. On the same time, the exhaust plume heats the nacelle in zone after exhaust pipe (see upper image in Fig. 15).

The recorded images had been used for determining the temperature distribution on the nacelle surface. The changes in nacelle surface temperature along the chosen reference lines in case of middle cruise regime are presented in Fig. 17.

Horizontal axis represents the length of line in pixels. Every point on the plot represents a pixel along the horizontal line,

from left end to the right end. P2 and P3 lines are going through

the exhaust tube where the temperatures are considered

inaccurate. However there is a well detectable zone with

[image:10.595.333.529.79.335.2]

heat-ing effect of exhaust gas on nacelle and propellers.

[image:10.595.341.528.378.503.2]

Fig. 15 Left side of left engine nacelle radiation with different temperature scales

[image:10.595.322.549.548.690.2]

Fig. 16 The exhaust pipe on left engine of aircraft AM-11 OKRA

Fig. 17 Changes in wall temperature along the chosen reference lines

The Fig. 18 demonstrates the real dependence of infrared radi-ation on engine installradi-ation. There are following zones of high temperature. Upper air intake louvres, which make visible hot

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parts of the engine. The exhaust pipe, with temperatures beyond the scale. The exhaust gases from the pipe heat the nose wheel

and ground around and after nose wheel to the temperatures

which seem to be extreme to rubber parts for longer period, as

well as for asphalt. Prolonged static operations of the engine set

for maximum continuous power during tests contributed to small chuckholes in the apron surface. And the latter, the exhaust from

oil cooler, located in the rear bottom of the cowling with tempera-tures within the scale. Because the applied temperature scale, the

temperature of the exhaust tube cannot be detected by use of this

image, in reality. However, the tube temperature is much greater (in analogy to Fig. 15), but additional temperature measurement

points located on the exhaust pipe and other points on the cowling

[image:11.595.319.528.46.179.2]

indicated, that the fumes are not so hot to damage the composite structure of the hood.

Fig. 18 Infrared images of nose section of ESPOSA aircraft demonstrator I-31 T

The Fig. 19 shows the thermal image of the aircraft I-31T noise section after 63 minutes ground test and 10 minutes cooling.

In Fig. 20 the temperature distribution shown at end of cruise phase along the chosen reference lines on nacelle wall of engine mounted in aircraft I-31T. This represents the best

possible temperature distribution during flight. Although the temperatures are possibly greater than during flight, but the distribution is as close as possible. The flight test program of

I-31T aeroplane included the measurements of temperature

distribution on the engine cowling during all phases of flight in order to satisfy the requirements of CS.23.1041 and follow -ing points concern-ing engine cool-ing. The points were located on the inner and outer surfaces of the cowling in areas of the

predicted high temeperature – exhaust traces, ECU, and gas

generator. The analysis revealed that the most critical points

on the engine cowling were located in the vicinity of exhaust

gas traces. The temperatures on the inner surfaces of upper

cowlings did not exceed 50 to 60 centigrade Celsius in gen -eral. However, recorded values reached temperatures up to 100

degrees Celsius in certain flight conditions. As well as may be

noticed in the Figs. 18 and 19, after engine shut down the tem-peratures in measuring points increased radically.

[image:11.595.304.552.224.387.2]

Fig. 19 Image after engine ground test and 10 minutes cooling (aircraft I-31 T)

Fig. 20 Changes in nacelle wall temperature along the chosen reference lines for engine built into aircraft I-31 T

8 Conclusions

Probably, the ESPOSA (2016) project was the first large

project developing and implementing a multi physical, namely aero and thermal optimization in integration of engine into the airframe. The goal was to minimize the aero-thermal losses for increasing the engines' effectivities and reducing the infrared radiations origin from engine - airframe structure. The study and optimization process were started by theoretical investiga-tions resulting to development of a CFD methods for aero-ther-mal integration of engine into airframe. After it a special prac-tical investigations were organized and realized.

The analysis of previous theoretical investigations and prac-tical measurements of the aircraft infrared radiations as well as

the theoretical results of the ESPOSA project allowed to define

the condition and methodology for infrared radiation mea-surements of the aircraft demonstrators. The paper introduced

the first systematic investigation of the small aircraft infrared

radiation. The full engine ground test was measured. The

mea-surements are fitted with theoretical investigations and show

[image:11.595.54.264.268.400.2]
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Acknowledgement

The work has been performed within the ESPOSA project (Grant Agreement No. ACP1-GA-2011-284859-ESPOSA)

funded by the 4th call of the FP7 Cooperation Work Programme.

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Figure

Fig. 1 Sources of aircraft infrared radiation (created by use of (Mahulikar et al., 2007)).
Fig. 2 CFM56 engine exhaust reflection on the wing bottom surface
Fig. 5 Infrared pictures of an engine nacelle at engine test
Fig. 6 Thermal radiation rays drawn from sensor point of view (a. (left side) - sensing the neighboring rays by steps of ∆θ and ∆ϕ,b
+5

References

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